CN104819056A - DSI air inlet with mix-compression profile surface and construction method of same - Google Patents
DSI air inlet with mix-compression profile surface and construction method of same Download PDFInfo
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Abstract
The invention discloses a DSI air inlet with a mix-compression profile surface, wherein the DSI air inlet successively includes a wedge-shaped shock wave compression surface, a first circular-conical shock wave compression surface, a second circular-conical shock wave compression surface, a third circular-conical shock wave compression surface, a fourth circular-conical shock wave compression surface and an inlet contour line of the air inlet. A plurality of air exhaust gaps are formed nearby an inlet throat. The invention also discloses a construction method of the DSI air inlet with a mix-compression profile surface. The DSI air inlet is improved in attack angle adaptability, is improved in compression shock wave stability, is improved in aerodynamic performance of an air vehicle and is enhanced in dynamic and steady-state performances of the air inlet.
Description
Technical field
The present invention relates to a kind of DSI intake duct, particularly relate to a kind of DSI intake duct mixing compression profile.
The invention still further relates to a kind of construction method mixing the DSI intake duct of compression profile.
Background technique
DSI intake duct (also known as without boundary layer diverter supersonic inlet, English name Diverterless Supersonic Inlet, is abbreviated as DSI intake duct) is that the one of rising in recent years can not mode supersonic inlet.Due to which eliminate supersonic inlet traditional remove boundary layer technological scheme every road, and the transverse-pressure gradient before changing dependence Fighter Inlet in compression profile orders about the technological scheme that boundary layer overflows to Fighter Inlet outward edge, therefore make supersonic inlet under the precondition keeping the air inlet performance parameters such as total pressure recovery coefficient substantially constant, simplify the structure, alleviate weight, reduce manufacture and working service cost; And due to the strong RCS signal source without boundary layer diverter, therefore also substantially improve the RCS Stealth Fighter of supersonic inlet.In view of the foregoing, DSI intake duct becomes with airbreathing motor the desirable air inlet system of the hypersonic vehicle of new generation being power plant.The course of new aircraft such as F35 and J10 remodeling are applied widely.
The DSI intake duct technological scheme of current use, usually take before Fighter Inlet, design the convex closure structure that is got rid of boundary layer, convex closure gets rid of boundary layer based on principle work series multistage circular cone compressing surface existing transverse-pressure gradient.The DSI intake duct technological scheme of series multistage circular cone compressing surface, there is when intake duct works in design point good air inlet performance, but for the aircraft that the flying condition such as flight attitude, speed excursion is large, when the flying condition change of aircraft causes intake duct to work in off-design point, to larger change be there is in the multishock produced by series multistage circular cone compressing surface before Fighter Inlet, thus causing inlet flow flowing state more to departing from away from design point, therefore intake duct air inlet performance declines more.
Therefore, need to provide a kind of new technological scheme to solve the problems referred to above.
Summary of the invention
For the aircraft that the flying condition such as flight attitude, speed excursion is large, need the DSI intake duct that novel, the technical issues that need to address of the present invention are to provide a kind of DSI intake duct mixing compression profile, make intake duct not only have good combination property at design point, and in the whole flight envelope of aircraft, still there is good combination property hold facility.
For solving technical problem of the present invention, the technical solution used in the present invention is:
A kind of DSI intake duct mixing compression profile, this DSI intake duct has wedge shape shock wave compression face, first Conical Shock Wave compressing surface, second Conical Shock Wave compressing surface, the 3rd road Conical Shock Wave compressing surface, the 4th road Conical Shock Wave compressing surface and Fighter Inlet profile line successively, is provided with many venting seams near described inlet throat.
Described venting Feng Wei tri-road.
Mix a construction method for the DSI intake duct of compression profile, it comprises the following steps:
1) the fuselage coordinates system of aircraft, is determined: before choosing aircraft, cusp is the longitudinal axis in fuselage coordinates system true origin, aircraft symmetry plane is Y-axis, vertical with Y-axis and the axle be directed upwardly is Z axis in aircraft symmetry plane, X-axis is determined according to right hand rule by Y-axis and Z axis;
2) reference datum in wedge shape shock wave compression face, is determined: chose Y-axis and " ∧ " shape ruled surface being symmetrical in YZ plane is the reference datum in wedge shape shock wave compression face;
3) Fighter Inlet profile line shape, is determined: according to the charge flow rate requirement of intake duct at design point, in conjunction with the appearance profile restriction of aircraft to intake duct and the reference datum position in wedge shape shock wave compression face, determine Fighter Inlet profile line shape (in certain plane vertical with wedge shape shock wave compression face reference datum draw direction), getting Fighter Inlet profile line is one section of circular arc line;
4) wedge shape shock wave compression face and Fighter Inlet profile line position, is determined: choose initial compression shock wave air-flow deflection angle δ 1, according to the Design of Inlet point angle of attack, freely flow Ma number and wedge shape shock wave compression face reference datum, determine the oblique shock angle β 1 of initial compression shock wave, thus determine Fighter Inlet profile line position, Y direction position is given; By Fighter Inlet profile line along the first oblique shock wave line direction projection in aircraft symmetry plane to wedge shape shock wave compression face reference datum, thus determine the initial edge line in wedge shape shock wave compression face; The initial edge line in wedge shape shock wave compression face is stretched along the first air-flow bend line in aircraft symmetry plane to Fighter Inlet place, produces wedge shape shock wave compression face;
5) the initial edge line of first Conical Shock Wave compressing surface, is determined: according to the relation M that total pressure recovery coefficient after multiple tracks oblique shock wave is maximum
a1sin β
1=M
a2sin β
2, β
1, M
a1refer to Angle of Shock Waves and the wavefront Ma number of front one shock wave respectively, β
2, M
a2refer to Angle of Shock Waves and the wavefront Ma number of rear one shock wave respectively, determine the oblique shock angle β 2 of first Conical Shock Wave; Then be half cone-apex angle with δ 1+ β 2, to be parallel to Y-axis and by the straight line in the arc profile line center of circle of Fighter Inlet for running shaft, with A, B 2 of Fighter Inlet edge line for base circumference passes through a little, do first Conical Shock Wave face, the intersection in first Conical Shock Wave face and wedge shape shock wave compression face is the initial edge line of first Conical Shock Wave compressing surface;
6) the auxiliary circle conical surface of first Conical Shock Wave compressing surface, is determined: according to Ma number after the wavefront Ma number of oblique shock angle β 2, first Conical Shock Wave and the ripple of wedge shape shock wave, determine the air-flow deflection angle δ 2 of first Conical Shock Wave compressing surface, with the straight line perpendicular to first Conical Shock Wave compressed lines on aircraft symmetry plane for bus, with the spin axis in first Conical Shock Wave face for running shaft, make the auxiliary circle conical surface of first Conical Shock Wave compressing surface;
7) first Conical Shock Wave compressing surface, is determined: projected with the auxiliary circle conical surface of face normal projection pattern to first Conical Shock Wave compressing surface by the initial edge line of first Conical Shock Wave compressing surface, produce the auxiliary projection line of first Conical Shock Wave compressing surface initial edge line, connect the corresponding points of first Conical Shock Wave compressing surface initial edge line and its auxiliary projection line to construct a curved surface, this curved surface is first Conical Shock Wave compressing surface;
8) the initial edge line of second Conical Shock Wave compressing surface, is determined: according to the relation M that total pressure recovery coefficient after multiple tracks oblique shock wave is maximum
a1sin β
1=M
a2sin β
2, β
1, M
a1refer to Angle of Shock Waves and the wavefront Ma number of front one shock wave respectively; β
2, M
a2refer to Angle of Shock Waves and the wavefront Ma number of rear one shock wave respectively, determine the oblique shock angle β 3 of second Conical Shock Wave; Then be half cone-apex angle with β 3+ δ 1+ δ 2, to be parallel to Y-axis and by the straight line in the arc profile line center of circle of Fighter Inlet for running shaft, with A, B 2 of Fighter Inlet edge line for base circumference passes through a little, do second Conical Shock Wave face, the intersection of second Conical Shock Wave face and first Conical Shock Wave compressing surface is the initial edge line of second Conical Shock Wave compressing surface;
9) the auxiliary circle conical surface of second Conical Shock Wave compressing surface, is determined: according to Ma number after the wavefront Ma number of oblique shock angle β 3, second Conical Shock Wave and the ripple of first Conical Shock Wave, determine the air-flow deflection angle δ 3 of second Conical Shock Wave compressing surface, with the straight line perpendicular to second Conical Shock Wave compressed lines on aircraft symmetry plane for bus, with the spin axis in second Conical Shock Wave face for running shaft, make the auxiliary circle conical surface of second Conical Shock Wave compressing surface;
10) second Conical Shock Wave compressing surface, is determined: projected with the auxiliary circle conical surface of face normal projection pattern to second Conical Shock Wave compressing surface by the initial edge line of second Conical Shock Wave compressing surface, produce the auxiliary projection line of second Conical Shock Wave compressing surface initial edge line, connect the corresponding points of second Conical Shock Wave compressing surface initial edge line and its auxiliary projection line to construct a curved surface, this curved surface is second Conical Shock Wave compressing surface;
11), follow-up Conical Shock Wave compressing surface is constructed: according to freely flowing the flow slowing down supercharging requirement before Ma number and DSI Fighter Inlet, determine the number of channels of circular cone pressure shock, and repeat the construction method of second Conical Shock Wave compressing surface, the follow-up Conical Shock Wave compressing surface of construction complete;
12) throat area of intake duct and corresponding profile, is determined: according to flying speed and the attitude scope of aircraft, determine the throat area of intake duct, and complete the quasi spline of inlet throat, inlet throat profile should with the extending section profile of smooth connection shock wave compression face and intake duct;
13) the extending section profile of intake duct, is determined; Determine the boundary layer floss hole position before Fighter Inlet and size; Determine the venting seam position near inlet throat and size: venting seam laterally lays multiple tracks along intake duct, its size take discharge quantity as the 2-3% design of intake duct charge flow rate.
Beneficial effect of the present invention: the DSI intake duct of mixing compression profile of the present invention, improves the angle of attack adaptability of DSI intake duct; Improve the pressure shock system stability of intake duct; Improve the aeroperformance of aircraft; Improve the dynamic of intake duct and steady-state behaviour.
Accompanying drawing explanation
Fig. 1 is system of coordinates definition schematic diagram.
Fig. 2 is the reference datum in wedge shape shock wave compression face.
Fig. 3 is Fighter Inlet profile line shape.
Fig. 4 is wedge shape shock wave compression face and Fighter Inlet profile line position.
Fig. 5 is the initial edge line of first Conical Shock Wave compressing surface.
Fig. 6 is the auxiliary circle conical surface of first Conical Shock Wave compressing surface.
Fig. 7 is first Conical Shock Wave compressing surface.
Fig. 8 is that boundary layer discharge seam affects schematic diagram to inlet characteristic.
Fig. 9 a is intake duct appearance schematic diagram.
Fig. 9 b be in Fig. 9 a A-A to schematic diagram.
Fig. 9 c is intake duct profile stereogram.
Fig. 9 d is I place's enlarged diagram in Fig. 9 b.
Fig. 9 e is II place's enlarged diagram in Fig. 9 c.
Figure 10 is intake duct shock wave compression face.
Figure 11 is intake duct partial sectional view (cutting plane is aircraft symmetry plane).
1, wedge shape shock wave compression face, 2, first Conical Shock Wave compressing surface, 3, second Conical Shock Wave compressing surface, the 4, the 3rd road Conical Shock Wave compressing surface, the 5, the 4th road Conical Shock Wave compressing surface, 6, Fighter Inlet profile line, 7, inlet throat, 8, venting seam.
Embodiment
Below in conjunction with the drawings and specific embodiments, the invention will be further described.Following examples, only for illustration of the present invention, are not used for limiting the scope of the invention.
Shown in Fig. 9 a, 9b, 9c, 9d, 9e, Figure 10, Figure 11, a kind of DSI intake duct mixing compression profile of the present invention, this DSI intake duct has wedge shape shock wave compression face 1, first Conical Shock Wave compressing surface 2, second Conical Shock Wave compressing surface the 3, the 3rd road Conical Shock Wave compressing surface the 4, the 4th road Conical Shock Wave compressing surface 5 and Fighter Inlet profile line 6 successively, is provided with three road venting seams 8 near inlet throat 7.
Mix a construction method for the DSI intake duct of compression profile, it comprises the following steps:
1) the fuselage coordinates system of aircraft, is determined: before choosing aircraft, cusp is the longitudinal axis in fuselage coordinates system true origin, aircraft symmetry plane is Y-axis, vertical with Y-axis and the axle be directed upwardly is Z axis in aircraft symmetry plane, X-axis is determined (as Fig. 1) according to right hand rule by Y-axis and Z axis;
2) reference datum in wedge shape shock wave compression face, is determined: chose Y-axis and " ∧ " shape ruled surface being symmetrical in YZ plane is the reference datum (as Fig. 2) in wedge shape shock wave compression face;
3) Fighter Inlet profile line shape, is determined: according to the charge flow rate requirement of intake duct at design point, in conjunction with the appearance profile restriction of aircraft to intake duct and the reference datum position in wedge shape shock wave compression face, determine Fighter Inlet profile line shape (in certain plane vertical with wedge shape shock wave compression face reference datum draw direction), getting Fighter Inlet profile line is one section of circular arc line (as Fig. 3);
4) wedge shape shock wave compression face and Fighter Inlet profile line position, is determined: choose initial compression shock wave air-flow deflection angle δ 1, according to the Design of Inlet point angle of attack, freely flow Ma number and wedge shape shock wave compression face reference datum, determine the oblique shock angle β 1 of initial compression shock wave, thus determine Fighter Inlet profile line position, Y direction position is given; By Fighter Inlet profile line along the first oblique shock wave line direction projection in aircraft symmetry plane to wedge shape shock wave compression face reference datum, thus determine the initial edge line in wedge shape shock wave compression face; The initial edge line in wedge shape shock wave compression face is stretched along the first air-flow bend line in aircraft symmetry plane to Fighter Inlet place, produces wedge shape shock wave compression face (as Fig. 4);
5) the initial edge line of first Conical Shock Wave compressing surface, is determined: according to the relation M that total pressure recovery coefficient after multiple tracks oblique shock wave is maximum
a1sin β
1=M
a2sin β
2, β
1, M
a1refer to Angle of Shock Waves and the wavefront Ma number of front one shock wave respectively, β
2, M
a2refer to Angle of Shock Waves and the wavefront Ma number of rear one shock wave respectively, determine the oblique shock angle β 2 of first Conical Shock Wave; Then be half cone-apex angle with δ 1+ β 2, to be parallel to Y-axis and by the straight line in the arc profile line center of circle of Fighter Inlet for running shaft, with A, B 2 of Fighter Inlet edge line for base circumference passes through a little, do first Conical Shock Wave face, the intersection in first Conical Shock Wave face and wedge shape shock wave compression face is the initial edge line (as Fig. 5) of first Conical Shock Wave compressing surface;
6) the auxiliary circle conical surface of first Conical Shock Wave compressing surface, is determined: according to Ma number after the wavefront Ma number of oblique shock angle β 2, first Conical Shock Wave and the ripple of wedge shape shock wave, determine the air-flow deflection angle δ 2 of first Conical Shock Wave compressing surface, with the straight line perpendicular to first Conical Shock Wave compressed lines on aircraft symmetry plane for bus, with the spin axis in first Conical Shock Wave face for running shaft, make the auxiliary circle conical surface (as Fig. 6) of first Conical Shock Wave compressing surface;
7) first Conical Shock Wave compressing surface, is determined: projected with the auxiliary circle conical surface of face normal projection pattern to first Conical Shock Wave compressing surface by the initial edge line of first Conical Shock Wave compressing surface, produce the auxiliary projection line of first Conical Shock Wave compressing surface initial edge line, connect the corresponding points of first Conical Shock Wave compressing surface initial edge line and its auxiliary projection line to construct a curved surface, this curved surface is first Conical Shock Wave compressing surface (as Fig. 7);
8) the initial edge line of second Conical Shock Wave compressing surface, is determined: according to the relation M that total pressure recovery coefficient after multiple tracks oblique shock wave is maximum
a1sin β
1=M
a2sin β
2, β
1, M
a1refer to Angle of Shock Waves and the wavefront Ma number of front one shock wave respectively; β
2, M
a2refer to Angle of Shock Waves and the wavefront Ma number of rear one shock wave respectively, determine the oblique shock angle β 3 of second Conical Shock Wave; Then be half cone-apex angle with β 3+ δ 1+ δ 2, to be parallel to Y-axis and by the straight line in the arc profile line center of circle of Fighter Inlet for running shaft, with A, B 2 of Fighter Inlet edge line for base circumference passes through a little, do second Conical Shock Wave face, the intersection of second Conical Shock Wave face and first Conical Shock Wave compressing surface is the initial edge line of second Conical Shock Wave compressing surface;
9) the auxiliary circle conical surface of second Conical Shock Wave compressing surface, is determined: according to Ma number after the wavefront Ma number of oblique shock angle β 3, second Conical Shock Wave and the ripple of first Conical Shock Wave, determine the air-flow deflection angle δ 3 of second Conical Shock Wave compressing surface, with the straight line perpendicular to second Conical Shock Wave compressed lines on aircraft symmetry plane for bus, with the spin axis in second Conical Shock Wave face for running shaft, make the auxiliary circle conical surface of second Conical Shock Wave compressing surface;
10) second Conical Shock Wave compressing surface, is determined: projected with the auxiliary circle conical surface of face normal projection pattern to second Conical Shock Wave compressing surface by the initial edge line of second Conical Shock Wave compressing surface, produce the auxiliary projection line of second Conical Shock Wave compressing surface initial edge line, connect the corresponding points of second Conical Shock Wave compressing surface initial edge line and its auxiliary projection line to construct a curved surface, this curved surface is second Conical Shock Wave compressing surface;
11), follow-up Conical Shock Wave compressing surface is constructed: according to freely flowing the flow slowing down supercharging requirement before Ma number and DSI Fighter Inlet, determine the number of channels of circular cone pressure shock, and repeat the construction method of second Conical Shock Wave compressing surface, the follow-up Conical Shock Wave compressing surface of construction complete;
12) throat area of intake duct and corresponding profile, is determined: according to flying speed and the attitude scope of aircraft, determine the throat area of intake duct, and complete the quasi spline of inlet throat, inlet throat profile should with the extending section profile of smooth connection shock wave compression face and intake duct;
13) the extending section profile of intake duct, is determined; Determine the boundary layer floss hole position before Fighter Inlet and size; Determine the venting seam position near inlet throat and size: venting seam laterally lays multiple tracks along intake duct, its size take discharge quantity as the 2-3% design of intake duct charge flow rate.
Advantage of the present invention is as follows:
1, improve the angle of attack adaptability of DSI intake duct: along with the progress of flying vehicles control technology, aircraft generally adopts BTT turning mode to improve the maneuverability of aircraft, therefore in the whole flight envelope of aircraft, the angle of sideslip of intake duct remains zero degree substantially, and the excursion of the intake duct angle of attack is larger, first order wedge shape shock wave compression face of the present invention, free incoming flow can be made to obtain initial compression while, the free incoming flow under any angle of attack can also be made to flow to change to keep the shock wave front flow direction in later stages conical pressure miniature face constant, thus improve the air inlet performance of DSI intake duct at off-design point, improve the angle of attack adaptability of DSI intake duct.
2, the pressure shock system stability of intake duct is improved: at design point or at off-design point, shock wave compression face, each road of the present invention edge line non-intersect (except the place of Fighter Inlet edge line two summit), shock wave compression face, each road is not intersected, thus avoid the flowing interference that shock interaction causes and be separated with boundary layer, improve the pressure shock system stability of intake duct in wide operating range.
3, improve the aeroperformance of aircraft: at design point, each road shock wave of intake duct can seal, thus reduces the wave resistance of intake duct, reduces the aerodynamical resistance of aircraft; Due to the flow insulated effect of the attached shock wave that first shock wave compression face (wedge shape shock wave compression face) produces, make the air-flow of aircraft lower surface be difficult to flow to upper surface, thus the aerodynamic lift of aircraft can be improved; The initial edge line in first shock wave compression face is specified by Fighter Inlet profile line, thus while the sealing of guarantee first shock wave three-dimensional omnidirectional, it also avoid the excessive reach of the aircraft pressure heart, ensure that the control stability of aircraft.
4, improve the dynamic of intake duct and steady-state behaviour: the present invention devises multiple tracks boundary layer discharge seam near inlet throat, when the fluctuation of inlet air flow parameter or engine chamber causes the fluctuation of air intake port back-pressure, the fluctuation of intake duct pressure perturbation is attached near boundary layer discharge seam by multiple tracks boundary layer discharge seam, be equivalent to the damping adding whole propulsion system, improve the antijamming capability of propulsion system, and keeping, under the condition that inlet stability margin is constant, improve the outlet total pressure recovery coefficient (as Fig. 8) of intake duct.
Claims (3)
1. one kind mixes the DSI intake duct of compression profile, it is characterized in that: this DSI intake duct has wedge shape shock wave compression face, first Conical Shock Wave compressing surface, second Conical Shock Wave compressing surface, the 3rd road Conical Shock Wave compressing surface, the 4th road Conical Shock Wave compressing surface and Fighter Inlet profile line successively, near described inlet throat, be provided with many venting seams.
2. a kind of DSI intake duct mixing compression profile according to claim 1, is characterized in that: described venting Feng Wei tri-road.
3. mix a construction method for the DSI intake duct of compression profile, it is characterized in that, it comprises the following steps:
1) the fuselage coordinates system of aircraft, is determined: before choosing aircraft, cusp is the longitudinal axis in fuselage coordinates system true origin, aircraft symmetry plane is Y-axis, vertical with Y-axis and the axle be directed upwardly is Z axis in aircraft symmetry plane, X-axis is determined according to right hand rule by Y-axis and Z axis;
2) reference datum in wedge shape shock wave compression face, is determined: chose Y-axis and " ∧ " shape ruled surface being symmetrical in YZ plane is the reference datum in wedge shape shock wave compression face;
3) Fighter Inlet profile line shape, is determined: according to the charge flow rate requirement of intake duct at design point, in conjunction with the appearance profile restriction of aircraft to intake duct and the reference datum position in wedge shape shock wave compression face, determine Fighter Inlet profile line shape, getting Fighter Inlet profile line is one section of circular arc line;
4) wedge shape shock wave compression face and Fighter Inlet profile line position, is determined: choose initial compression shock wave air-flow deflection angle δ 1, according to the Design of Inlet point angle of attack, freely flow Ma number and wedge shape shock wave compression face reference datum, determine the oblique shock angle β 1 of initial compression shock wave, thus determine Fighter Inlet profile line position, Y direction position is given; By Fighter Inlet profile line along the first oblique shock wave line direction projection in aircraft symmetry plane to wedge shape shock wave compression face reference datum, thus determine the initial edge line in wedge shape shock wave compression face; The initial edge line in wedge shape shock wave compression face is stretched along the first air-flow bend line in aircraft symmetry plane to Fighter Inlet place, produces wedge shape shock wave compression face;
5) the initial edge line of first Conical Shock Wave compressing surface, is determined: according to the relation M that total pressure recovery coefficient after multiple tracks oblique shock wave is maximum
a1sin β
1=M
a2sin β
2, β
1, M
a1refer to Angle of Shock Waves and the wavefront Ma number of front one shock wave respectively, β
2, M
a2refer to Angle of Shock Waves and the wavefront Ma number of rear one shock wave respectively, determine the oblique shock angle β 2 of first Conical Shock Wave; Then be half cone-apex angle with δ 1+ β 2, to be parallel to Y-axis and by the straight line in the arc profile line center of circle of Fighter Inlet for running shaft, with A, B 2 of Fighter Inlet edge line for base circumference passes through a little, do first Conical Shock Wave face, the intersection in first Conical Shock Wave face and wedge shape shock wave compression face is the initial edge line of first Conical Shock Wave compressing surface;
6) the auxiliary circle conical surface of first Conical Shock Wave compressing surface, is determined: according to Ma number after the wavefront Ma number of oblique shock angle β 2, first Conical Shock Wave and the ripple of wedge shape shock wave, determine the air-flow deflection angle δ 2 of first Conical Shock Wave compressing surface, with the straight line perpendicular to first Conical Shock Wave compressed lines on aircraft symmetry plane for bus, with the spin axis in first Conical Shock Wave face for running shaft, make the auxiliary circle conical surface of first Conical Shock Wave compressing surface;
7) first Conical Shock Wave compressing surface, is determined: projected with the auxiliary circle conical surface of face normal projection pattern to first Conical Shock Wave compressing surface by the initial edge line of first Conical Shock Wave compressing surface, produce the auxiliary projection line of first Conical Shock Wave compressing surface initial edge line, connect the corresponding points of first Conical Shock Wave compressing surface initial edge line and its auxiliary projection line to construct a curved surface, this curved surface is first Conical Shock Wave compressing surface;
8) the initial edge line of second Conical Shock Wave compressing surface, is determined: according to the relation M that total pressure recovery coefficient after multiple tracks oblique shock wave is maximum
a1sin β
1=M
a2sin β
2, β
1, M
a1refer to Angle of Shock Waves and the wavefront Ma number of front one shock wave respectively; β
2, M
a2refer to Angle of Shock Waves and the wavefront Ma number of rear one shock wave respectively, determine the oblique shock angle β 3 of second Conical Shock Wave; Then be half cone-apex angle with β 3+ δ 1+ δ 2, to be parallel to Y-axis and by the straight line in the arc profile line center of circle of Fighter Inlet for running shaft, with A, B 2 of Fighter Inlet edge line for base circumference passes through a little, do second Conical Shock Wave face, the intersection of second Conical Shock Wave face and first Conical Shock Wave compressing surface is the initial edge line of second Conical Shock Wave compressing surface;
9) the auxiliary circle conical surface of second Conical Shock Wave compressing surface, is determined: according to Ma number after the wavefront Ma number of oblique shock angle β 3, second Conical Shock Wave and the ripple of first Conical Shock Wave, determine the air-flow deflection angle δ 3 of second Conical Shock Wave compressing surface, with the straight line perpendicular to second Conical Shock Wave compressed lines on aircraft symmetry plane for bus, with the spin axis in second Conical Shock Wave face for running shaft, make the auxiliary circle conical surface of second Conical Shock Wave compressing surface;
10) second Conical Shock Wave compressing surface, is determined: projected with the auxiliary circle conical surface of face normal projection pattern to second Conical Shock Wave compressing surface by the initial edge line of second Conical Shock Wave compressing surface, produce the auxiliary projection line of second Conical Shock Wave compressing surface initial edge line, connect the corresponding points of second Conical Shock Wave compressing surface initial edge line and its auxiliary projection line to construct a curved surface, this curved surface is second Conical Shock Wave compressing surface;
11), follow-up Conical Shock Wave compressing surface is constructed: according to freely flowing the flow slowing down supercharging requirement before Ma number and DSI Fighter Inlet, determine the number of channels of circular cone pressure shock, and repeat the construction method of second Conical Shock Wave compressing surface, the follow-up Conical Shock Wave compressing surface of construction complete;
12) throat area of intake duct and corresponding profile, is determined: according to flying speed and the attitude scope of aircraft, determine the throat area of intake duct, and complete the quasi spline of inlet throat, inlet throat profile should with the extending section profile of smooth connection shock wave compression face and intake duct;
13) the extending section profile of intake duct, is determined; Determine the boundary layer floss hole position before Fighter Inlet and size; Determine the venting seam position near inlet throat and size: venting seam laterally lays multiple tracks along intake duct, its size take discharge quantity as the 2-3% design of intake duct charge flow rate.
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
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| CN115653754A (en) * | 2022-12-12 | 2023-01-31 | 中国航空工业集团公司西安飞机设计研究所 | Supersonic air inlet system with three wave systems for fixing compression surface |
Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101029597A (en) * | 2007-03-22 | 2007-09-05 | 南京航空航天大学 | Fixed geometrical supersonic-speed and high supersonic-speed adjusting air inlet |
| CN102230848A (en) * | 2007-12-03 | 2011-11-02 | 哈尔滨工业大学 | Method for acquiring optimal classification surface of hypersonic speed air inlet starting/non-starting mode |
| CN103950543A (en) * | 2014-04-18 | 2014-07-30 | 南京航空航天大学 | Aircraft supersonic air inlet channel with variable deflation system |
| CN203819494U (en) * | 2014-05-23 | 2014-09-10 | 厦门大学 | Waverider and inlet integrated device controllable in lateral pressure gradient |
| CN204755085U (en) * | 2015-05-05 | 2015-11-11 | 江西洪都航空工业集团有限责任公司 | DSI intake duct that mixes compression profile |
-
2015
- 2015-05-05 CN CN201510225032.6A patent/CN104819056B/en active Active
Patent Citations (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101029597A (en) * | 2007-03-22 | 2007-09-05 | 南京航空航天大学 | Fixed geometrical supersonic-speed and high supersonic-speed adjusting air inlet |
| CN102230848A (en) * | 2007-12-03 | 2011-11-02 | 哈尔滨工业大学 | Method for acquiring optimal classification surface of hypersonic speed air inlet starting/non-starting mode |
| CN103950543A (en) * | 2014-04-18 | 2014-07-30 | 南京航空航天大学 | Aircraft supersonic air inlet channel with variable deflation system |
| CN203819494U (en) * | 2014-05-23 | 2014-09-10 | 厦门大学 | Waverider and inlet integrated device controllable in lateral pressure gradient |
| CN204755085U (en) * | 2015-05-05 | 2015-11-11 | 江西洪都航空工业集团有限责任公司 | DSI intake duct that mixes compression profile |
Non-Patent Citations (2)
| Title |
|---|
| 李淑艳: "高超声速进气道设计及性能研究", 《中国优秀硕士论文全文数据库》 * |
| 杨应凯: "枭龙飞机Bump进气道设计", 《南京航空航天大学学报》 * |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115653754A (en) * | 2022-12-12 | 2023-01-31 | 中国航空工业集团公司西安飞机设计研究所 | Supersonic air inlet system with three wave systems for fixing compression surface |
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